1. Field of the Invention
The present invention relates to a combustion gas supply control device.
2. Description of Related Art
In operating a missile, high mobility at the terminal guidance is required. One approach is to use a multi-pulse rocket motor and activate an appropriate pulse at the terminal guidance to reaccelerate the missile. For example, in operating a missile equipped with a two-pulse rocket motor, the missile is flown toward a target by being accelerated by a first pulse (combustion of a first propellant) and then the missile is reaccelerated by a second pulse (combustion of a second propellant) when coming close to the target, to enhance high mobility at the terminal guidance. Therefore, when aiming at a remote target, it is necessary to increase the amount of the first propellant consumed in the first pulse in order to increase either burn time or thrust, or both the burn time and thrust.
However, in a case of a two-pulse rocket motor disclosed in Patent Document 1 (Japanese Patent No. 3231778) and Patent Document 2 (Japanese Patent Publication JP-2005-171970), the first propellant and the second propellant are arranged in an axial direction of the rocket motor and also there is a limit to a length of the rocket motor in terms of equipment or storage. Therefore, the first propellant cannot have enough length. In this case, it is highly possible that an initial burning area becomes small and thus necessary initial thrust cannot be obtained.
Moreover, as for a barrier membrane and a barrier membrane holder of the two-pulse rocket motor disclosed in Patent Document 1 and Patent Document 2, breakability and durability of the barrier membrane at the time when a second igniter operates are unclear. When a barrier membrane is broken at an unintended position, combustion of the second propellant or a combustion gas flow is disturbed. In some cases, a nozzle may be blocked up with the broken barrier membrane.
In a case where the first igniter and the second igniter are arranged in series in the axial direction of the rocket motor disclosed in Patent Document 2, the igniters have cantilever long and thin structure. Therefore, the structure needs to be strengthened in order to secure the strength of the igniters against such environment as vibration at the time of operation. This causes increase in structural weight, which is unsuitable for a long rocket motor.
The inventors of the present application have proposed, in Patent Document 3 (Japanese Patent No. 4719182), a two-pulse rocket motor which can solve the problems of the techniques disclosed in Patent Document 1 and Patent Document 2. The two-pulse rocket motor disclosed in Patent Document 3 will be described below.
FIG. 1 is a longitudinal sectional view showing an example of the two-pulse rocket motor disclosed in Patent Document 3. FIG. 2 is a sectional view taken along a line A-A in FIG. 1.
The two-pulse rocket motor generates two-step thrust by first combusting a first propellant 4 and then, after a certain period of time has passed, combusting a second propellant 5. Therefore, the second propellant 5, until starting to be burned, needs to withstand high-temperature combustion gas and high pressure generated as a result of combustion of the first propellant 4.
As shown in FIGS. 1 and 2, a nozzle 2 having an exhaust hole 12 at the center for exhausting the combustion gas is fixed to a rear portion of a cylindrical motor case 1. A motor head 3 is fixed to a front portion of the motor case 1, and a first igniter 6 for combusting the first propellant 4 is fixed to the motor head 3.
The first propellant 4 and the second propellant 5 both in a hollow tubular shape (i.e. an internal-burning type propellant shape or an internal-end-burning type propellant shape) are loaded within the motor case 1. The second propellant 5 is arranged on an outer periphery of a front portion of the first propellant 4. It should be noted that the shape of the first propellant 4 and the second propellant 5 each may be a hollow cylinder, a hollow tube with a polygonal outer surface and/or a polygonal inner surface, or a hollow cone.
The first propellant 4 and the second propellant 5 are isolated from each other by a barrier membrane 10. A highly heat-resistant rubber such as EPDM rubber, silicone rubber, silicone rubber or EPDM rubber containing such inorganic fiber as Kevlar fiber can be used as the barrier membrane 10.
A second igniter 8 for combusting the second propellant 5 is provided at a forward end of the second propellant 5.
An operation of the two-pulse rocket motor shown in FIGS. 1 and 2 is as follows. The first igniter 6 starts operating in response to an external signal and hence the first propellant 4 starts burning (combusting). At this point of time, the barrier membrane 10 is not exposed to high-temperature combustion gas. After that, when the first propellant 4 has been combusted to the position of the barrier membrane 10, the barrier membrane 10 is exposed to high-temperature combustion gas. After a certain period of time has passed from completion of combustion of the first propellant 4, the second igniter 8 starts operating in response to an external signal and hence the second propellant 5 starts burning (combusting).
According to the two-pulse rocket motor shown in FIGS. 1 and 2 as described above, an inner surface of the first propellant 4 is exposed to a burning region 11 over almost the full length of the motor case 1 in the axial direction, and thereby an initial burning area can be secured. Therefore, there is no need to provide the inner surface of the first propellant 4 with a large slit.
Moreover, since the second propellant 5 is arranged on the outer periphery of the first propellant 4, a burning area of the second propellant 5 does not become extremely smaller than a burning area of the first propellant 4. Therefore, the nozzle 2 can be shared by the first propellant 4 and the second propellant 5.
In addition, since the second propellant 5 is arranged on the outer periphery of the first propellant 4 and the barrier membrane 10 is provided between the first propellant 4 and the second propellant 5, a time during which the barrier membrane 10 is exposed to the high-temperature combustion gas can be shortened as much as possible. In other words, heat protection of the barrier membrane 10 is achieved.
Furthermore, heat protection of the second igniter 8 is achieved by the barrier membrane 10 as in the case of the second propellant 5. The second igniter 8 is burned down due to its operation as expected.
It should be noted that an initial burning surface of the second propellant 5 means a surface which burns from an initial stage when the second propellant 5 starts burning, namely, a surface which is first exposed to the burning region 11 when the second propellant 5 starts burning. In the case of the example shown in FIG. 1, the initial burning surface of the second propellant 5 includes a cylindrical inner surface and a ring-shaped rear surface of the second propellant 5.
FIG. 3 is a longitudinal sectional view showing another example of the two-pulse rocket motor disclosed in Patent Document 3. FIG. 4 is a sectional view taken along a line B-B in FIG. 3.
In the example shown in FIGS. 3 and 4, the barrier membrane 10 is divided into two parts to provide a weak section (joint section). More specifically, the barrier membrane 10 covering the second propellant 5 comprises: an aft barrier membrane 10a in a circular truncated cone shape placed on a rear surface of the second propellant 5; and an inner barrier membrane 10b in a tubular shape placed on the inner surface of the second propellant 5. Respective ends of the aft barrier membrane 10a and the inner barrier membrane 10b are bonded with each other by a fire-resistant adhesive over an entire periphery, to serve as the weak section. The weak section (joint section) is not broken during combustion of the first propellant 4 but is certainly broken by pressure of gas generated by the operation of the second igniter 8 or combustion of the second propellant 5.
FIG. 5 is a longitudinal sectional view showing a deformation state of the barrier membrane 10 at the time of combustion of the second propellant 5. FIG. 6A is a sectional view taken along a line C-C in FIG. 5. FIG. 6B is a sectional view taken along a line D-D in FIG. 5.
At the time of combustion of the second propellant 5, the inner barrier membrane 10b, which is a large part of the barrier membrane 10, is deformed toward the center of the motor case 1 and is held at the forward portion of the motor case 1 where a combustion gas flow of the second propellant 5 is relatively slow. Meanwhile, a break portion of the aft barrier membrane 10a is deformed to be turned up backward along the combustion gas flow. Therefore, such an effect as breakability and durability of the barrier membrane 10 (the aft barrier membrane 10a and the inner barrier membrane 10b) become clear and certain and can be obtained in addition to the above-mentioned effects.
It should be noted that the same effects as in the case of the above-mentioned divided structure can be obtained even when the barrier membrane 10 is formed integrally and a cutoff line or a notch is provided at a position to be broken.
The second igniter 8 may be placed at a rearward end surface of the second propellant 5 that is closer to the weak section (joint section) of the barrier membrane 10 (the aft barrier membrane 10a and the inner barrier membrane 10b). In this case, certainty of breakage of the barrier membrane 10 (the aft barrier membrane 10a and the inner barrier membrane 10b) becomes higher.
In the case of the two-pulse rocket motor as described above, the second igniter 8 and the first igniter 6 are independent of each other across the barrier membrane 10, in terms of structure. Therefore, it is possible to secure the strength of the igniters against such environment as vibration at the time of operation, without strengthening the structure of the igniters to increase structural weight even in a case of a long motor.